A gas turbine engine typically includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure compressor includes annular arrays (“rows”) of vanes that direct air entering the engine into downstream, rotating blades of the compressor. Collectively one row of compressor vanes and one row of compressor blades make up a “stage” of the compressor. Similarly, the high pressure turbine includes annular rows of vanes that direct the gases exiting the combustor into downstream, rotating blades of the turbine. Collectively one row of nozzle vanes and one row of turbine rotor blades make up a “stage” of the turbine. Typically, both the compressor and turbine include a plurality of successive stages.
In operation, a hot gas having a temperature between approximately 3000 and approximately 3500 degrees Fahrenheit flows through the turbine and each of its successive stages. However, the high temperatures experienced by the turbine during operation may stress the components of the turbine, specifically the turbine rotor blades. As such, in an effort to cool the turbine rotor blades, bleed air from the compressor flows through a cooling passage defined within the turbine rotor blade. The cooling passage generally extends from a root portion of the turbine rotor blade to a blade tip of the turbine rotor blade along a radial direction. Further, when the bleed air exits the cooling passage through outlets formed on the turbine rotor blade, the bleed air mixes with the hot gas. Thus, the bleed air may not be used to cool other components within the turbine.
Accordingly, a system and method for cooling components of a gas turbine engine would be welcomed within the technology. In particular, a system and method that more effectively uses a cooling airflow would be particularly beneficial.